Method of orientation of space vehicle and device for its implementation

FIELD: physics, navigation.

SUBSTANCE: group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections.

EFFECT: improvement of orientation accuracy and operational reliability in case of failures of orientation angle sensor and sensor of angular velocity of space vehicle rotation.

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The invention relates to the field of rocket technology, namely the management of non-stationary objects spacecraft.

The known method the orientation of the spacecraft, consisting in the measurement signal of the orientation angle and the angular velocity signal, the formation of the reference signal and the signal processing control spacecraft [1].

A device for implementing the method of the orientation of the spacecraft containing the serially connected first adder, a first amplifier, a second adder, the engine flywheel, the third adder, the spacecraft, the first and second outputs of which are connected respectively to the input of the sensor of angular velocity and entry angle sensor, the output of the second amplifier is connected with the second input of the second adder [1].

The known method the orientation of the spacecraft and a device for implementing the method have the disadvantage that is low-precision orientation and low reliability due to failure of the sensor orientation angle and the angular rate sensor of the rotation of the spacecraft.

To eliminate these disadvantages of the known method, the orientation of the spacecraft and device for its implementation the proposed method differs in that form signal estimate of the orientation angle and the signal evaluation corner is karasti rotation of the spacecraft, form a reference signal of the orientation angle and the reference angular velocity signal form the signal evaluation management, determine the signal of the difference signal of the orientation angle and signal evaluation of the orientation angle, determine the difference signals of the angular velocity signal and the signal estimate angular velocity, determine the signal of the difference signal of the orientation angle and the reference signal of the orientation angle, determine the difference signals of the angular velocity signal and the reference signal of the angular velocity, determine the signal difference signal control and signal evaluation management and determine the correction signal of the reference signal, the correction signal of the reference signal of the mathematical model and the assessment of the external signal interference by the formulas respectively

$$U_{K_0}={\lambda }_0\left({\stackrel{}{\epsilon }}_0+a_1^m{\epsilon }_0+a_0^m{\displaystyle \underset{0}{\overset{t}{\int }}{\epsilon }_{0}dt}\right)$$,

$$U_K=\lambda \left(\stackrel{}{\epsilon }+a_1\epsilon +a_0\right)$$,

$${\stackrel{}{M}}_{\mathrm{in}}=K_m{\displaystyle \underset{0}{\overset{t}{\int }}\Delta Udt}$$,

where λ₀, λ,a₀,a₁=const>0, K_m- gain, ε₀signal to the difference signal of the orientation angle and the reference signal of the orientation angle, ε is the signal of the difference signal of the orientation angle and signal evaluation of the orientation angle, ΔU is the difference signals signal control and signal evaluation control, which corrects the signal evaluation of the orientation angle and the signal estimation of angular velocity and use them for orientation of the spacecraft, and the device for its realization, characterized in that it additionally contains a reference model of the primary circuit orientation, memory block, nine adders, six amplifiers, five integrators, two normally-open switch, five normally-closed switches, the output of the fourth adder connected in series through the third amplifier, the fifth adder, the sixth adder, the first model of the engine flywheel, the first integrato is, the second integrator, the seventh adder, the first normally-closed switch, the third integrator, the eighth adder, the second normally-closed switch and a memory unit connected to the first input of the fourth adder, the output of the sensor of angular velocity through the third normally-closed switch connected to the input of the second amplifier, and through a series-connected ninth adder and the fourth amplifier to the second input of the eighth adder, a third input connected to the input of the third integrator, the output of the first integrator is connected to a second input of the ninth adder, and through the first normally-open switch - over input of the second amplifier, the output of the fifth amplifier is connected to a second input of the fifth adder, the output of the second integrator is connected to a second input of the fourth adder, and a second normally-open switch with a first input of the first adder, a second input connected to the third input of the fourth adder, the output of the sensor orientation angle is connected with the second input of the seventh adder and a fourth normally-closed switch with the first input of the first adder, the output of the second adder connected in series through the tenth adder, fifth normally-closed switch and the fourth integrator connected to the second input of the Estai adder, the output of which is connected to a second input of the tenth adder, the second input of the first adder connected in series through a reference model of the primary circuit orientation, the eleventh adder, the fifth integrator and the sixth amplifier connected to the fourth input of the fourth adder, the second output of the reference model of the primary circuit orientation connected in series through the twelfth and seventh adder amplifier is connected to the fifth input of the fourth adder, the output of the eleventh adder through eighth amplifier is connected to the sixth input of the fourth adder, and through the ninth amplifier with a third input of the first adder, the output of the fifth integrator through the tenth amplifier connected to the fourth input of the first adder, the input of the second amplifier is connected to a second input of the twelfth adder whose output through the eleventh amplifier is connected to the fifth input of the first adder.

The essence of the proposed method and device illustrated in Fig.1, which depicts a block diagram of the device, and Fig.2 - structural diagram of the standard model of the primary circuit orientation.

In the drawings, the following notation:

1 - the fourth adder;

2 - the third amplifier;

3 - the fifth adder;

4 - model of the engine-flywheel;

5 - the sixth adder;

6 - the fifth amplifier;

7 is basic to the tour orientation (EYE) of the spacecraft;

8 - the first integrator;

9 - the second integrator;

10 - second normally-open switch;

11 - the first normally-open switch;

12 - second normally-closed switch;

13 is a block of memory;

14 - fourth integrator;

15 is a mathematical model of the EYE;

16 - the eighth adder;

17 - third integrator;

18 - the seventh adder;

19 - the fourth amplifier;

20 - ninth adder;

21 - tenth adder;

22 - the fifth normally-closed switch;

23 - the first adder;

24 - the first amplifier;

25 - the second adder;

26 - engine-flywheel;

27 - third adder;

28 - spacecraft;

29 - angular rate sensor;

30 - second amplifier;

31 - the third normally-closed switch;

32 fourth normally-closed switch;

33 - angle sensor orientation;

34 - the first normally-closed switch;

35 - reference model EYE;

36 - fifth integrator;

37 - the eleventh adder;

38 - twelfth adder;

39 - sixth amplifier;

40 - seventh amplifier;

41 - eighth amplifier;

42 - the eleventh amp;

43 - ninth amplifier;

44 - tenth amplifier;

45, 46 respectively of the eleventh and twelfth amp;

47, 48, the first and second outputs of the reference model EYE, respectively;

50 - trine is dcity adder;

51 - thirteenth amp;

52 - fourteenth adder;

53 - reference model engine flywheel;

54 sixth integrator;

55 - seventh integrator.

Operates the device to implement the method of the orientation of the spacecraft in the following way. The reference signal φ_C(t) is supplied simultaneously to the main control circuit 7 and a mathematical model of the EYE 15 (see Fig.1).

As can be seen from Fig.1, the EYE 7 are connected in series, the first adder 23, the first amplifier 24, a second adder 25, the engine-flywheel 26, the third adder 27, the spacecraft 28 and the angular rate sensor 29. The output of the sensor orientation angle 33 through the fourth normally-closed switch 32 is connected to the first adder 23, forming a negative feedback on the orientation angle φ(t), and from the output of the angular rate sensor 29 angular velocity$$\stackrel{}{\phi }\left(t\right)$$is fed to the input of the second adder 25, forming a negative feedback on the angular velocity$$\stackrel{}{\phi }\left(t\right)$$. To the second input of the third adder 27 receives external interference M_in(t). The parameters of the spacecraft 28 operates a mul is applicationa disturbance F(t).

A mathematical model of the EYE 15 is composed of the same EYE 7 of the circuit elements, connected to the fourth adder 1, the third amplifier 2, the fifth adder 3, the sixth adder 5, the model of the engine-flywheel 4 and the model of the spacecraft, made in the form of series-connected first integrator 8 and the second integrator 9. In the mathematical model of the EYE 15 also shows a negative feedback on the evaluation of the angular velocity$$\stackrel{}{\stackrel{}{\phi }}\left(t\right)$$and evaluation of the orientation angle$$\stackrel{}{\phi }\left(t\right)$$with outputs of the first integrator 8 to the input of the fifth adder 3 and the output of the second integrator 9 to the input of the fourth adder 1.

Due to the fact that the spacecraft 28 are external interference M_in(t) and F(t), and the mathematical model of the EYE 15 of the spacecraft 28 external noise F(t) and M_in(t) do not apply, the evaluation of$$\stackrel{}{\stackrel{}{\phi }}\left(t\right)$$and$$\stackrel{}{\phi }\left(t\right)$$ will not coincide with their real values, respectively$$\stackrel{}{\phi }\left(t\right)$$and φ(t).

Therefore, the control U(t) in an EYE 7 at the output of the second adder 25 is different from the control U_m(t) in the model EYE 15 at the output of the sixth adder 5. We will simulate the effect of external interference M_inin the EYE 7 in the form of assessment external interference$${\stackrel{}{M}}_{\mathrm{in}}$$in the model of the EYE 15.

With this purpose, using the tenth adder 21 fifth normally closed switch 22 and the fourth integrator 14 will form a value$${\stackrel{}{M}}_{\mathrm{in}}$$as

$${\stackrel{}{M}}_{\mathrm{in}}=K_{gm}\left(U-U_m\right)$$,

where K_gmtransfer function model of the engine-flywheel 4. Thus it is enough to adjust the value of U_m(t) at the output of the sixth adder 5. In addition, the moment of inertia J(t) of the spacecraft in the General case differs from the evaluation of $$\stackrel{}{J}$$. To determine the true value of moment of inertia J(t) will form using the component method of formation control based on the use of Lyapunov functions, the signal adjustment ΔJ(t), which, as shown in [2], will be

$$\Delta J\left(t\right)=\lambda {\phi }_C\left(\stackrel{}{\epsilon }+a_1\epsilon +a_0{\displaystyle \underset{0}{\overset{t}{\int }}\epsilon dt}\right)$$,

which corresponds to the correction signal U_toevaluation of angle$$\stackrel{}{\phi }\left(t\right)$$and angular velocity$$\stackrel{}{\stackrel{}{\phi }}\left(t\right)$$equal to

,

where λ,and₀,and₁=const>0.

To ensure the sustainability of the orientation system comprising a mathematical model of the EYE 15 and the EYE 7 requires a certain sign of the correction signal U_to(t). The sign of the signal U_to(t) defines only the component of$$\stackrel{}{\epsilon }+a_1\epsilon +{\int }_0^t\epsilon dt$$because the value of$${\phi }_C^2$$always greater than or equal to zero.

Therefore, using the rule of signs in the construction of adaptive control systems component according to the method of formation of management [2], we can take the value of the correction signal U_to(t) in the form

$$U_K\left(t\right)=\lambda \left(\stackrel{}{\epsilon }+a_1\epsilon +a_0{\displaystyle \underset{0}{\overset{t}{\int }}\epsilon dt}\right)$$ .

Due to the fact that U_to(t) depends on the integral of the error ε(t), it is obvious that the output of the third integrator 17 is continuously changed up until the input is not a value of zero. This means that when ε(t)=0 are equal

$$\phi \left(t\right)=\stackrel{}{\phi }\left(t\right)$$and$$\stackrel{}{\phi }\left(t\right)=\stackrel{}{\stackrel{}{\phi }}\left(t\right)$$.

To implement the correction signal used in the seventh adder 18, the first normally-closed switch 34, the third integrator 17, the eighth adder 16, the ninth adder 20, the fourth amplifier 19, the second normally-closed switch 12 and the memory unit 13, the output of which is connected to the third input of the fourth adder (see Fig.1).

According to the proposed method the orientation of the reference signal φ_C(t) is the input signal for the main circuit orientation 7 and the mathematical model of the EYE 15. The control signal U(t) is generated at the output of the second adder 25, and a signal evaluation control U_m(t) at the output of the sixth adder 5. The outputs of the sensor of angular skorosti and sensor orientation angle 33 are formed, respectively, the angular velocity of $$\stackrel{}{\phi }\left(t\right)$$rotation of the spacecraft 28 and the orientation angle φ(t).

The signal increment ΔJ(t) values of moment of inertia J(t) spacecraft 28 is a function of the error ε(t):

$$\epsilon \left(t\right)=\phi \left(t\right)-\stackrel{}{\phi }\left(t\right)$$

and the reference signal φ_C(t).

The signal evaluation of the external moment (interference)$${\stackrel{}{M}}_{\mathrm{in}}\left(t\right)$$is formed by using a tenth adder 21, the fifth normally closed switch 22, the fourth integrator 14, the sixth adder 5 and model of the engine-flywheel 4.

You can now use the mathematical model of the EYE 15 as angular rate sensor 29 and the angle sensor 33, if at the same time to open the normally-closed switches 22, 37, 34, 31 and 32 and to close the normally-open switches 10 and 11.

A mathematical model of the EYE 15 it is reasonable to use as speed sensors 29 and orientation angle 33 to until the error ε(t) is within the PPRs is of valid values, determined, obviously, by changes in the external interference M_in(t) and values of moment of inertia J(t) spacecraft 28.

Thus, the connection of the EYE 7 and the mathematical model of the EYE 15 allows you to generate the correction signal U_to(t)for a mathematical model of the EYE 15 and the signal ΔU(t), simulating the effect of signal interference in the form of moment M_in(t).

As a result, the outputs of φ and$$\stackrel{}{\phi }$$on the outputs, respectively, the EYE 7 and the mathematical model of the EYE 15 at the end of the transition process are the same. This allows the use of a mathematical model of the EYE as the identifier of the parameter J(t) and phase coordinates φ(t) and$$\stackrel{}{\phi }\left(t\right)$$.

However, in this case, the EYE 7 runs with error ε₀(t)=φ-φ_mwhere φ is the output of the EYE 7, and φ_mthe output of the reference model EYE 35.

The output φ(t) contains a static error ε₁(t), as described in [1]. The error caused by the perturbing moment M_in(t), which is input to the reference model EYE 35 is not received.

The formation of the correction signal$$U_{K_1}\left(t\right)$$

$$U_{K_0}\left(t\right)={\lambda }_0\left({\stackrel{}{\epsilon }}_0+a_1^m{\epsilon }_0+a_0^m{\displaystyle \underset{0}{\overset{t}{\int }}{\epsilon }_{0}dt}\right)$$,

where λ₀,$$a_1^m$$,$$a_0^m=const>0$$determine the quality of the transient process and the error of the device orientation.

When this signal$$U_{K_0}\left(t\right)$$corrected the reference signals as the EYE 7 and the mathematical model of the EYE 15 in order to compensate for the error ε₀(t).

Implemented a reference model EYE (shown in Fig.2) in the form of connection of circuit elements 45, 46, 50, 51, 52, 53, 54 and 55 copies of the EYE 7 under the condition that the transfer function of the angular rate sensor 29 and sensor orientation angle 33 b is izki to the unit.

Implementation of the correction signal$$U_{K_0}\left(t\right)$$by using circuit elements depicted in Fig.1: 35÷42.

Reference model EYE 35 operates similarly to the EYE 7. The difference is that it doesn't hindrance M_B(t).

The formation of the correction signal$$U_{K_0}\left(t\right)$$for the EYE 7 and the mathematical model of the EYE 15 and its implementation is carried out similarly to the correction of the mathematical model of the EYE 15 signal U_to(t).

Use in the way that the orientation of the spacecraft and device for its implementation of channel estimation$$\stackrel{}{\phi }\left(t\right)$$and$$\stackrel{}{\stackrel{}{\phi }}\left(t\right)$$- mathematical model of the EYE 15 - allows you to obtain the technical effect is to increase the accuracy of the orientation and reliability of operation in case of failure of sensor orientation angle 33 and the angular rate sensor 29 of rotation of the spacecraft 28.

Literature

1. Vasiliev Century. N. Attitude control system for spacecraft / C. N. Vasiliev. - M.: FSUE NPP VNIIEM", 2009. S. 149-156 (prototype).

2. Losev A. J. Method for the synthesis of adaptive control systems with reference model. Devices and systems. Management, monitoring, diagnostics. 2007. No. 1. C. 2-6.

1. The way the orientation of the spacecraft, consisting in the measurement signal of the orientation angle and the angular velocity signal, the formation of the reference signal and the signal processing control spacecraft, wherein forming the signal evaluation of the orientation angle and the signal evaluation of the angular velocity of rotation of the spacecraft, form the reference signal of the orientation angle and the reference angular velocity signal form the signal evaluation management, determine the signal of the difference signal of the orientation angle and signal evaluation of the orientation angle, determine the difference signals of the angular velocity signal and the signal estimate angular velocity, determine the signal of the difference signal of the orientation angle and the reference signal of the orientation angle, determine the difference signals of the angular velocity signal and the reference signal of the angular velocity, determine the signal difference signal control and signal evaluation management is to determine the correction signal of the reference signal, the correction signal of the reference signal of the mathematical model and the assessment of the external signal interference by the formulas accordingly
,
,
,
where λ₀, λ,a₀,a₁= const > 0, K_m- gain, ε₀signal to the difference signal of the orientation angle and the reference signal of the orientation angle, ε is the signal of the difference signal of the orientation angle and signal evaluation of the orientation angle, ΔU is the difference signals signal control and signal evaluation control, which corrects the signal evaluation of the orientation angle and the signal estimation of angular velocity and use them for orientation of the spacecraft.

2. Device orientation of the spacecraft containing the serially connected first adder, a first amplifier, a second adder, the engine flywheel, the third adder, the spacecraft, the first and second outputs of which are connected respectively to the input angular rate sensor and the sensor input of the orientation angle, the output of the second amplifier is connected with the second input of the second adder, characterized in that it additionally contains a reference model of the primary circuit orientation, memory block, nine adders, six amplifiers, five integrators, two normally-open switch, five normally causes the output switches, the output of the fourth adder connected in series through the third amplifier, the fifth adder, the sixth adder, the first model of the engine flywheel, the first integrator, the second integrator, the seventh adder, the first normally-closed switch, the third integrator, the eighth adder, the second normally-closed switch and a memory unit connected to the first input of the fourth adder, the output of the sensor of angular velocity through the third normally-closed switch connected to the input of the second amplifier, and through a series-connected ninth adder and the fourth amplifier to the second input of the eighth adder, a third input connected to the input of the third integrator, the output of the first integrator is connected to a second input of the ninth adder, and through the first normally-open switch - over input of the second amplifier, the output of the fifth amplifier is connected to a second input of the fifth adder, the output of the second integrator is connected to a second input of the fourth adder, and a second normally-open switch with a first input of the first adder, a second input connected to the third input of the fourth adder, the output of the sensor orientation angle is connected with the second input of the seventh adder and a fourth normally-closed switch with the first input of the first adder, in the course of the second adder connected in series through the tenth adder, fifth normally-closed switch and the fourth integrator is connected to a second input of the sixth adder, the output of which is connected to a second input of the tenth adder, the second input of the first adder connected in series through a reference model of the primary circuit orientation, the eleventh adder, the fifth integrator and the sixth amplifier connected to the fourth input of the fourth adder, the second output of the reference model of the primary circuit orientation connected in series through the twelfth and seventh adder amplifier is connected to the fifth input of the fourth adder, the output of the eleventh adder through eighth amplifier is connected to the sixth input of the fourth adder, and through the ninth amplifier with a third input of the first adder, the fifth output of the integrator through the tenth amplifier connected to the fourth input of the first adder, the input of the second amplifier is connected to a second input of the twelfth adder whose output through the eleventh amplifier is connected to the fifth input of the first adder.